It is known that in general jet engines operate at higher efficiency if they operate at higher temperatures. If the operating temperature of a jet engine can be increased by 100.degree. the efficiency of operation of the engine can be significantly improved. Jet engines last more than 10 years in service. If the fuel consumed by a jet engine is reduced by a significant degree over the 10 of more years of expected life of a jet engine, then there is a cost saving in the operation of engine which is very substantial and which permits the engine to be formed at higher costs. The higher engine cost is more than offset by the lower costs of operation of the engine.
The operation of jet engines at higher temperatures also results in a greater thrust-to-weight ratio. In other words, if the same jet engine design is maintained but the materials are altered so that the temperature of operation of the engine is increased then the net result will be that the engine will be found to have a higher thrust-to-weight ratio than the same engine operated at the lower temperature. The materials which are employed in a jet engine which is operated at higher temperature must have greater temperature capability. Alternatively, if materials can be found which operate at the same or higher temperature but which have lower density then a higher thrust-to-weight ratio may be achieved. Further, it is possible to design engines which have material with greater temperature capability and with lower density and this combination also yields engines with greater thrust-to-weight ratios.
Not all of the portions of a jet engine are operated at the same temperature. The portions of the engine which operate at the highest temperature presently operate below 2200.degree. F. The present invention contemplates the modification of the components in the hottest sections of the engines, and particularly of the coatings on the component elements of the hottest sections, so that the component temperature in these sections will operate at temperatures above 2400.degree. F. These temperatures are far greater than encountered in present components. Most materials, such as nickel base alloys, which are presently employed in jet engines are molten at temperatures above 2450-2500.degree. F.
Various metallic systems have been investigated for the hottest components of jet engines to determine the maximum temperature at which they may be employed. The lower density, but lower ductility, ceramic systems are competing with the metallic systems for applications in the hottest components of jet engines. Some of the metallic systems which have been considered include metal matrix composites in which a strengthening component such as a filament is incorporated within a metal matrix. Also low density intermediate phases and intermetallic compounds have been considered for such high temperature applications.
One of the problems which has been associated with the development of metallic systems for high stress capability at high temperatures is that of oxidation of the metallic component at the high temperatures. The choice of metals which can be employed is broadened by the availability of a coating, such as is provided pursuant to the present invention, which will withstand the engine environment.
Presently the nickel base alloys are protected by an alumina-forming metallic coating. Such a coating has a sufficient Al reservoir in the coating to re-form the protective scale when spallation of the oxide from the outer surface occurs. Present iron, cobalt, and nickel base alloys and their alumina-forming metallic coatings are intended for use at lower temperatures below their melting points. The nickel base alloys are not the most reactive metals and in cases where the protective coating is lost the nickel alloy can withstand the engine environment in its uncoated condition for relatively short periods so that loss of the coating for such short periods is not catastrophic to engine performance.
However for a refractory metal or intermetallic system which operates at service temperatures of greater than 2200.degree. F., once a breach of a protective coating is formed the substrate metal may be degraded very rapidly either by oxidation loss of metal cross-section or by environmental embrittlement. For composite systems having a reinforcement element embedded within a matrix metal designed for service at temperatures greater than 2200.degree. F., the large surface area between the matrix and the reinforcement may serve as a rapid diffusion path for such oxidation and/or embrittlement. Accordingly, the demands on a coating and the requirements for a coating on a component to protect the component from the engine environment is much more severe than is the case for the components formed of the nickel base alloys which operate at lower temperatures. One such requirement is that a coating have the capability of rapidly healing of any breach of the protective oxide due to spallation or similar cause so that a "fail-safe" performance of the base metal and coating system may be achieved.